1. Technical Field
The present invention relates to rotary blade assemblies for gas turbine engines in general, and to rotary blade tip structures in particular.
2. Background Information
Gas turbine engines are typically designed to minimize the clearance between rotary blade tips and the surrounding casing. Such designs decrease air leakage between the blade tips and the casing, and thereby improve the efficiency of the engine. Some prior art designs include an abradable material disposed on an interior portion of the casing surrounding the rotary blade tips. During initial use of such prior art designs, the rotary blades extend radially outward and engage the abradable material, creating a trench within the abradable material. During subsequent use of such prior art designs, the rotary blade tips extend into the trench and thereby create a decreased leakage air path between the rotary blade tips and the abradable seal. These designs work reasonably well, but can also have drawbacks relating to mechanical durability of the blade tips. For example, prior art rotary blades made of durable materials (i.e., materials sufficiently durable to prevent blade tip failure) often make the rotary blade undesirably heavy. Another prior art attempt to solve blade tip durability involved anodizing rotary blade tips to strengthen them. This approach can be problematic because the anodizing can cause cracking in the fan blade.
What is needed, therefore, are rotary blades for use in gas turbine engines, which blades are sufficiently durable so as to prevent the blade tips from being damaged if they engage an abradable seal material, and which rotary blades overcome the problems discussed above.